Fan drive gear system

ABSTRACT

A gas turbine engine includes a fan section and a speed change mechanism for driving the fan section. The speed change mechanism is an epicyclic gear train. A torque frame surrounds the speed change mechanism and includes a plurality of fingers. A bearing support is attached to the plurality of fingers. A first fan section support bearing is mounted forward of the speed change mechanism and a second fan section bearing is mounted on the bearing support aft of the speed change mechanism. The second fan section bearing is a fan thrust bearing.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is continuation on U.S. application Ser. No. 14/950,326filed Nov. 24, 2015, which is a CIP of U.S. application Ser. No.14/848,978 filed Sep. 9, 2015 which claims priority to U.S. ProvisionalApplication No. 62/054,506 which was filed on Sep. 24, 2014. Thisapplication also claims priority to U.S. Provisional Application No.62/085,924 which was filed on Dec. 1, 2014 through U.S. application Ser.No. 14/950,326.

BACKGROUND

Turbomachines, such as gas turbine engines, typically include a fansection, a turbine section, a compressor section, and a combustorsection. Turbomachines may employ a geared architecture connecting thefan section and the turbine section. The compressor section typicallyincludes at least a high-pressure compressor and a low-pressurecompressor. The compressors include rotors that rotate separately from arotor of fan. To maximize performance of such turbomachines, variousrecent engine architectures have been proposed in which the fan rotatesin a first direction and at a first speed as compared to a low pressurecompressor which rotates in the opposite direction and at a higherspeed. These recent engine architectures can also be improved.

SUMMARY

In one exemplary embodiment, a gas turbine engine includes a fan sectionand a speed change mechanism for driving the fan section. The speedchange mechanism is an epicyclic gear train. A torque frame surroundsthe speed change mechanism and includes a plurality of fingers. Abearing support is attached to the plurality of fingers. A first fansection support bearing is mounted forward of the speed change mechanismand a second fan section bearing is mounted on the bearing support aftof the speed change mechanism. The second fan section bearing is a fanthrust bearing.

In a further embodiment of any of the above, the fan thrust bearingengages a gas turbine static structure and the bearing support.

In a further embodiment of any of the above, the bearing support isattached to distal ends of the plurality of fingers.

In a further embodiment of any of the above, the epicyclic gear train isa planetary gear system that includes a sun gear in communication with afan drive turbine and a planet carrier in communication with the fansection.

In a further embodiment of any of the above, the torque frame includes afirst end for engaging the fan section and second end supporting thesecond fan section bearing.

In a further embodiment of any of the above, each of the plurality offingers include at least one groove.

In a further embodiment of any of the above, the bearing supportincludes a plurality of tangs that engage a corresponding one of the atleast one groove.

In a further embodiment of any of the above, at least one groove islocated on a radially inner side of a corresponding one of the pluralityof fingers.

In a further embodiment of any of the above, the speed change mechanismis at least partially axially aligned with a compressor section.

In a further embodiment of any of the above, there is a high pressurecompressor with a compression ratio of approximately 20:1 or greater anda fan bypass ratio of approximately 10 or greater.

In a further embodiment of any of the above, there is a low speed spool,an intermediate spool, and a high speed spool.

In another exemplary embodiment, a speed change mechanism for a gasturbine engine includes a planetary gear system. A torque framesurrounds the speed change mechanism. The torque frame includes aplurality of fingers. A bearing support is attached to a downstream endof the plurality of fingers for supporting a fan section supportbearing.

In a further embodiment of any of the above, the planetary gear systemincludes a sun gear that is in communication with a fan drive turbine. Aplanet carrier is in communication with the fan section.

In a further embodiment of any of the above, the plurality of fingersengage grooves in a planet carrier of the speed change mechanism. Thebearing support is attached to a distal end of the plurality of fingers.

In a further embodiment of any of the above, a fan thrust bearing isattached to the bearing support.

In another exemplary embodiment, a method of assembling a gas turbineengine includes supporting a fan section on a first fan section supportbearing located forward of a speed change mechanism. The speed changemechanism is a planetary gear system supported by a torque frame thathas a plurality of fingers attached to a bearing support. The fansection is supported on a second fan section support bearing attached tothe bearing support located aft of the speed change mechanism.

In a further embodiment of any of the above, the torque frame includes afirst end for engaging the fan section and a second end attached to abearing support for supporting the second fan section support bearing.

In a further embodiment of any of the above, the bearing supportincludes a plurality of tangs that each engages a groove on acorresponding one of the plurality of fingers.

In a further embodiment of any of the above, the gas turbine engineincludes a low speed spool, an intermediate spool, and a high speedspool. A low pressure compressor includes at least one compressor stageand no more than five compressor stages.

In a further embodiment of any of the above, a fan section and a lowpressure compressor is supported on the first fan section supportbearing located forward of the speed change mechanism.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is an enlarged schematic view of a portion of the example gasturbine engine of FIG. 1.

FIG. 3 is a sectional view taken along line 3-3 of FIG. 2.

FIG. 4 is a sectional view taken along line 4-4 of FIG. 3.

FIG. 5 is a perspective view of a portion of a speed change mechanism.

FIG. 6 is a schematic view of another example gas turbine engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26, and a turbine section 28. Alternative enginesmight include an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42 through an input shaft 41, a first (or low)pressure compressor 44 and a first (or low) pressure turbine 46. Theinner shaft 40 is connected to the fan 42 through a speed changemechanism, which in exemplary gas turbine engine 20 is illustrated as ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a second (or high) pressure compressor 52 and a second (orhigh) pressure turbine 54. The second pressure compressor 52 includes acompression ratio of approximately 20:1 or greater. A combustor 56 isarranged in exemplary gas turbine 20 between the high pressurecompressor 52 and the high pressure turbine 54. A mid-turbine frame 57of the engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The mid-turbineframe 57 further supports bearing systems 38 in the turbine section 28.The inner shaft 40 and the outer shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A which iscollinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft(10,668 meters), with the engine at its best fuel consumption—also knownas “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is theindustry standard parameter of 1 bm of fuel being burned divided by 1 bfof thrust the engine produces at that minimum point. “Low fan pressureratio” is the pressure ratio across the fan blade alone, without a FanExit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosedherein according to one non-limiting embodiment is less than about 1.45.“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram°R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150ft/second (350.5 meters/second).

As shown in FIG. 1, the low pressure compressor 44 is axially alignedwith the geared architecture 48 so that fan exit guide vanes 58 arelocated further aft to reduce noise from the gas turbine engine 20.

As shown in FIG. 2, the geared architecture 48 includes a sun gear 60that is mounted to the input shaft 41 that is attached to the innershaft 40. Accordingly, the sun gear 60 is driven by the input shaft 41.Surrounding the sun gear 60 is a plurality of planet gears 62 that aresupported on bearings 63 by a planet carrier 64. The planet gears 62 aresurrounded by a ring gear 66 that is mounted to the engine staticstructure 36 with a flexible coupling 68 that allows the gearedarchitecture 48 to flex to allow for proper alignment between thevarious elements of the geared architecture 48 during operation. Theplanet carrier 64 is attached to a fan drive shaft 70 through a torqueframe 72.

As shown in FIGS. 2 and 3, an oil transfer bearing 74 directs lubricantfrom a plurality of stationary oil tubes 76, such as a first stationaryoil tube 80, a second station oil tube 82, and a third stationary oiltube 84, into the fan drive shaft 70 and the torque frame 72 tolubricate the geared architecture 48. The oil transfer bearing 74includes a plurality of inputs to provide lubricant to those portions ofthe geared architecture 48 that require lubrication during operation.For example, oil from the first stationary oil tube 80 is intended tolubricate at least one of the bearing systems 38, oil from the secondstationary oil tube 82 is intended to lubricate the bearings 63 in thegeared architecture 48, and oil from the third stationary oil tube 84 isintended to lubricate the sun gear 60, planet gears 62, and the ringgear 66 of the geared architecture 48. Though three oil tube inputs areshown herein, other numbers of oil tubes are contemplated herein.

FIG. 4 illustrates a sectional view of the oil transfer bearing 74. Inthe illustrated example, the oil transfer bearing 74 includes a firstrace 86, a second race 88, and a third race 90 each having a rectangularshape that extend around an interior surface 92 of a stationary bearing74 a.

A first oil conduit 94 extends axially through the fan drive shaft 70and is in communication with the first race 86 via a first opening 96. Asecond oil conduit 98 extends axially through the fan drive shaft 70 andis in communication with the second race 88 via a second opening 100. Athird oil conduit 102 extends axially through the fan drive shaft 70 andis in communication with the third race 90 via a third opening 104.

As the fan drive shaft 70 and the rotating bearing 74 b rotate withinthe stationary bearing 74 a, the first, second, and third openings 96,100, 104 are constantly in alignment with the first, second, and thirdraces 86, 88, 90, respectively. This allows oil to flow across arotating gap between the stationary bearing 74 a and the rotatingbearing 74 b through the first, second, and third openings 96, 100, 104to the first, second, and third oil conduits 94, 98, 102, respectively,to provide lubrication to the necessary areas in the gas turbine engine20.

As shown in FIG. 2, the geared architecture 48 is located axiallybetween a fan roller bearing 106 and a fan thrust bearing 108. The fanroller bearing 106 engages the fan drive shaft 70 and the staticstructure 36. The fan thrust bearing 108 engages the static structure 36and a support ring 110 attached to the torque frame 72.

As shown in FIG. 5, the torque frame 72 includes a base ring 112 thatattaches to the fan drive shaft 70 on a first axial side and a pluralityof fingers 114 that extend from a second axial side of the base ring112. The plurality of fingers 114 slideably engage corresponding grooves126 in the planet carrier 64. At least one torque frame pin opening 118extends through each of the plurality of fingers 114 and aligns with atleast one carrier pin opening 120 in the planet carrier 64 to accept apin 122 that will pass through the at least one torque frame pinopenings 118 and into that at least one carrier pin openings 120 to lockthe torque frame 72 from moving axially relative to the planet carrier64.

As shown in FIG. 2, the plurality of fingers 114 extend beyond an axialdownstream side of the planet carrier 64. In the illustrated example,the axial downstream ends of the plurality of fingers 114 includelocking slots 128 on radially inner sides of each of the plurality offingers 114 to engage a support ring 130. The locking slots 128 extendbetween opposing sides of each of the plurality of fingers 114.

The support ring 130 includes a plurality of tangs 132 extending outwardfrom a radially outer side of the support ring 130. The plurality oftangs 132 are circumferentially spaced around an outer perimeter of thesupport ring 130 to align with the locking slots 128 of each of theplurality of fingers 114.

In order to attach the support ring 130 to the torque frame 72, thegrooves 126 on the planet carrier 64 are circumferentially aligned withthe plurality of fingers 114. Then the planet carrier 64 is movedaxially toward the torque frame 72 until the torque frame pin openings118 align with the carrier pin openings 120. The pins 122 then extendthrough the torque frame pin openings 118 and the carrier pin openings120 to lock the planet carrier 64 relative to the torque frame 72 fromrelative axial movement. The plurality of fingers 114 engaging thegrooves 126 prevents the torque frame 72 from rotating relative to theplanet carrier 64 and allows torque to be transferred from the planetcarrier 64 into the torque frame 72 and then through the fan drive shaft70.

Once the planet carrier 64 is secured relative to the torque frame 72,the support ring 130 is aligned so that the plurality of tangs 132 arecircumferentially aligned with open areas circumferentially locatedbetween the plurality of fingers 114. The support ring 130 then movesaxially towards the torque frame 72 until the plurality of tangs 132 areaxially aligned with the locking slots 128. The locking ring 130 is thenrotated either clockwise or counterclockwise until the plurality oftangs 132 are aligned with the plurality of fingers 114 and locatedwithin the locking slots 128. A lock nut 134 is treaded onto a threadedportion of the support ring 130 to prevent the support ring 130 fromrotating relative to the torque frame 72.

Once the support ring 130 has been secured to the torque frame 72, thesupport ring 130 will then support a radially inner side of the fanthrust bearing 108. This allows for a compact packaging of the gearedarchitecture 48 that can reduce the overall length of the gas turbineengine 20 and allow the geared architecture to be straddled by both thefan roller bearing 106 and the fan thrust bearing 108.

FIG. 6 illustrates another example gas turbine engine 20′. The examplegas turbine engine 20′ is similar to the gas turbine engine 20 exceptwhere described below or shown in the Figures. The gas turbine engine20′ is disclosed herein as a three-spool turbofan that generallyincorporates a fan section 22′, a compressor section 24′, a combustorsection 26′, and a turbine section 28′.

The exemplary engine 20′ generally includes a low speed spool 30′, anintermediate spool 31, and a high speed spool 32′ mounted for rotationabout an engine central longitudinal axis A relative to an engine staticstructure 36 via several bearing systems 38.

The low speed spool 30′ generally includes an inner shaft 40′ thatinterconnects a fan 42 through an input shaft 41′, a first (or low)pressure compressor 44′ and a first (or low) pressure turbine 46′. Theinner shaft 40′ is connected to the fan 42 through a speed changemechanism, which in the exemplary gas turbine engine 20′ is illustratedas a geared architecture 48′ to drive the fan 42 at a lower speed thanthe low speed spool 30′. In one example, the low pressure compressor 44′includes at least one compressor stage and no more than five compressorstages. In another example, the low pressure compressor 44′ includes atleast two compressor stages and no more than four compressor stages.

The intermediate spool 31′ includes an intermediate shaft 43 thatinterconnects a third (or intermediate) pressure compressor 51 with athird (or intermediate) pressure turbine 53.

The core airflow is compressed by the low pressure compressor 44′, theintermediate pressure compressor 51, and the high pressure compressor52′, mixed and burned with fuel in the combustor 56′, then expanded overthe high pressure turbine 54′, the intermediate pressure turbine 53, andlow pressure turbine 46′. The mid-turbine frame 57 includes airfoils 59which are in the core airflow path C. The turbines 46′, 53, and 54′rotationally drive the respective low speed spool 30′, intermediatespool 31, and the high speed spool 32′ in response to the expansion. Itwill be appreciated that each of the positions of the fan section 22′,compressor section 24′, combustor section 26′, turbine section 28′, andgeared architecture 48′ may be varied. For example, the gearedarchitecture 48 may be located aft of combustor section 26′ or even aftof turbine section 28′, and fan section 22′ may be positioned forward oraft of the location of the geared architecture 48′.

Although the different non-limiting embodiments are illustrated ashaving specific components, the embodiments of this disclosure are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed and illustrated in these exemplary embodiments,other arrangements could also benefit from the teachings of thisdisclosure.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldunderstand that certain modifications could come within the scope ofthis disclosure. For these reasons, the following claim should bestudied to determine the true scope and content of this disclosure.

What is claimed is:
 1. A gas turbine engine comprising: a fan section; aspeed change mechanism for driving the fan section, wherein the speedchange mechanism is an epicyclic gear train; a torque frame surroundsthe speed change mechanism and includes a plurality of fingers, whereindistal ends of the plurality of fingers are located downstream of aplanet carrier of the epicyclic gear train; a bearing support attachedto the plurality of fingers; and a first fan section support bearingmounted forward of the speed change mechanism and a second fan sectionsupport bearing mounted on the bearing support aft of the speed changemechanism, wherein the second fan section support bearing is a fanthrust bearing.
 2. The gas turbine engine of claim 1, wherein the fanthrust bearing engages a gas turbine static structure on a radiallyouter side and the fan thrust bearing engages the bearing support on aradially inner side with the radially outer side and the radially innerside relative to a central longitudinal axis of the gas turbine engine.3. The gas turbine engine of claim 2, wherein the epicyclic gear trainis a planetary gear system including a sun gear in communication with afan drive turbine and a planet carrier in communication with the fansection and the bearing support is attached to distal ends of theplurality of fingers and is axially spaced from the planet carrier. 4.The gas turbine engine of claim 3, wherein the torque frame includes afirst end for engaging the fan section and second end supporting thesecond fan section bearing.
 5. The gas turbine engine of claim 4,wherein each of the plurality of fingers include at least one groove. 6.The gas turbine engine of claim 5, wherein the bearing support includesa plurality of tangs that engage a corresponding one of the at least onegroove.
 7. The gas turbine engine of claim 6, wherein the at least onegroove is located on a radially inner side of a corresponding one of theplurality of fingers.
 8. The gas turbine engine of claim 1, wherein thespeed change mechanism is at least partially axially aligned with acompressor section.
 9. The gas turbine engine of claim 1, furthercomprising a high pressure compressor with a compression ratio ofapproximately 20:1 or greater, a fan bypass ratio of approximately 10 orgreater, a low speed spool, an intermediate spool, and a high speedspool.
 10. A speed change mechanism for a gas turbine engine comprising:a planetary gear system including a planet carrier supporting aplurality of planet gears; a torque frame surrounding the speed changemechanism, wherein the torque frame includes a plurality of fingers witha distal end of the plurality of fingers axially downstream of theplanet carrier; and a bearing support forming a ring and axially spacedfrom the planet carrier, the bearing support is attached to a downstreamend of the plurality of fingers for supporting a fan section supportbearing.
 11. The speed change mechanism of claim 10, wherein theplanetary gear system includes a sun gear in communication with a fandrive turbine and the planet carrier is in communication with the fansection and the plurality of fingers engage grooves in radiallyoutermost surface of the planet carrier of the speed change mechanism.12. The speed change mechanism of claim 10, wherein the plurality offingers engage grooves in the planet carrier of the speed changemechanism, a distal end of each of the plurality of fingers includes aradially inward facing groove for accepting a tang on the bearingsupport and the bearing support forms a ring.
 13. The speed changemechanism of claim 12, further comprising a fan thrust bearing having aradially inner side attached to the bearing support and configured torotate at the same speed and in the same direction as the bearingsupport.
 14. A method of assembling a gas turbine engine comprising:supporting a fan section on a first fan section support bearing locatedforward of a speed change mechanism wherein the speed change mechanismis a planetary gear system supported by a torque frame having aplurality of fingers attached to a bearing support, wherein theplanetary gear system includes a planet carrier and the bearing supportis axially spaced from the planet carrier and is rotatable relative tothe plurality of fingers and the planet carrier; and supporting the fansection on a second fan section support bearing attached to the bearingsupport located aft of the speed change mechanism.
 15. The method ofclaim 14, wherein the torque frame includes a first end for engaging thefan section and a second end attached to the bearing support forsupporting the second fan section support bearing.
 16. The method ofclaim 15, wherein the bearing support includes a plurality of tangs thateach engages a groove on a corresponding one of the plurality offingers.
 17. The method of claim 16, wherein the gas turbine engineincludes a low speed spool, an intermediate spool, and a high speedspool and a low pressure compressor including at least one compressorstage and no more than five compressor stages.
 18. The method of claim14 further comprising supporting a fan section and a low pressurecompressor on the first fan section support bearing located forward ofthe speed change mechanism.
 19. The gas turbine engine of claim 1,wherein each of the plurality of fingers include at least one groove andthe bearing support includes a plurality of tangs that engage acorresponding one of the at least one groove.